The present invention relates to a rotary wing aircraft, for example of the helicopter type, in which a rotor, fitted with blades and constituting said rotary wing, is pitch-controlled entirely electrically.
It is known that standard helicopters comprise numerous mechanical and hydromechanical members required for the control of the collective pitch and of the cyclic pitch of the blades of their rotor. Such members are expensive and moreover exhibit considerable mass.
There have already been attempts to reduce the mass and the cost of such a helicopter (see for example patents U.S. Pat. No. 5,387,083, U.S. 5,409,183 and U.S. 5,626,312) by providing electric flight controls which operate the deflection of flaps, each of them being mounted on one of said blades of the rotor and being able to aerodynamically control the pitch of the associated blade, thus making it possible to do away with a few mechanical and hydromechanical members such as jacks, pumps and hydraulic circuits, but none of the other mechanical members.
Furthermore, in these known embodiments of helicopters with electric controls and pitch-control flaps, instability problems arise which are such that the standard aeronautic safety criteria may not be satisfied with regard to elementary or combined breakdowns which are not highly improbable.
The object of the present invention is to remedy these drawbacks. It relates to an aircraft in which the changes of pitch of the blades are obtained with the aid of such flaps, which is enhanced in such a way as to make it possible to do away with almost all the mechanical and hydromechanical members for controlling the pitch of the rotor, while ensuring perfect safety of use.
To this end, according to the invention, the aircraft comprises:
a fuselage;
control members available to a pilot, which are disposed in said fuselage and are able to produce flying instructions;
at least one rotor revolving with respect to said fuselage and catering for the functions of uplift of said aircraft and of displacement of the latter according to roll and pitching axes as well as in vertical and horizontal translation, said rotor comprising at least two blades whose pitch can be controlled so as to allow said rotor to cater for said functions;
orientable flaps carried by said blades so as to control their pitch, each of said blades comprising a plurality of flaps aligned along the span of said blade and each blade comprising as many flaps, arranged identically, as any one of the other blades;
electric actuation devices carried by said blades so as to actuate said orientable flaps with a view to the control of the pitch of said blades;
a flight control device disposed in said fuselage and producing, from said flying instructions and from signals representative of flight parameters, a plurality of control commands for the plurality of said electric actuation devices;
a plurality of links provided between said flight control device and each of said electric actuation devices so as to address to each of these electric devices, through said interface, one of said control commands; and
a stationary/rotary interface between said fuselage and said rotor allowing said plurality of links to transmit said control commands to said electric actuators,
xe2x80x83the aircraft in which:
said flight control devices comprises as many distinct control paths as there are flaps on each blade;
said interface comprises as many distinct transmission paths as there are distinct control paths;
said distinct control paths of said flight control device are respectively associated with said distinct transmission paths of said interface so as to form, with said links, as many distinct control channels as there are flaps on each blade; and
each of said control channels produces and conveys the control commands for all the electric actuation devices of the flaps of the same rank on said blades.
Thus, by virtue of the plurality of flaps mounted on each of said blades and of the plurality of control commands addressed, electrically and/or optically, to the electric actuation devices of said flaps, it is possible to effect all of the pitchwise controls of the rotor of said aircraft electrically. It is therefore possible to do away with all the mechanical and hydromechanical members used as standard for this purpose, namely the pitch levers, the control linkages, the sliding and rotating swashplates, the pumps and hydraulic circuits, the jacks, the control rods and levers and the combiner.
This results in a considerable lightening and appreciable simplification of the mechanical assemblies specific to a rotary wing aircraft and in particular to a helicopter. Moreover, such a rotor control system is easily transposable and reconfigurable from one type of aircraft to another, the gist of the modifications then consisting in a new setting of the various parameters with a limited impact on the hardware devices.
Moreover, in the aircraft according to the present invention, the multiplicity of links between the flight control device and the flaps of the blades makes it possible to obtain a multiplicity of independent control channels, having no common member.
Thus, in the event of breakdown of hardware or software elements, the aircraft according to the invention can continue to be flown in complete safety, although possibly with reduced performance. Flight safety is therefore ensured.
The flight control device can be of the type described in American patent U.S. Pat. No. 6,059,225.
To further increase reliability and safety, said control channels may be embodied according to different technologies. It is then advantageous for the control channels of flaps close to the free ends of said bladesxe2x80x94these being the flaps which exhibit the greatest aerodynamic effectivenessxe2x80x94to be embodied according to the technology regarded as the most reliable.
Preferably, each electric actuation device is housed in the corresponding blade, in immediate proximity to the flap with which it is associated. Thus, the mechanical link between an electric actuation device and its flap can be short, so that the control of said flap is immediate and direct and that long control rods routed through the blade can be avoided.
For additional reasons of safety, each electric actuation device comprises at least two electric actuatorsxe2x80x94jack or motorxe2x80x94mounted in parallel.
The flaps for controlling the pitch of the blades could be mounted, on the latter, on their leading edge side. However, for reasons of stability it is preferable for said flaps to be arranged in the trailing edge of said blades.
Moreover, in order to cope with any failure of a control channel through a defect in one of its elements specific to a single blade, which would run the risk of a considerable imbalance of the forces exerted on the rotor, there is provided an autosurveillance system permanently verifying, for each control channel, the conformity of the implementation of the deflections of the flaps of the channel with regard to the control commands. In the event of nonconformity of implementation, all the flaps of the channel are mechanically immobilized. The positional holding of the flaps which is necessary to prevent any aerodynamic or aeroelastic flutter phenomenon, must be effective whatever failure is at issue, including cases of loss of electrical power for the relevant channel.
Furthermore, by virtue of the fact that, in the rotary wing aircraft according to the invention, all the pitch control commands are electrical or optical, auxiliary signals can easily be appended to said commands so as to implement additional functions, such as, as will be described hereinafter, an antiresonance function, a wing autoadjustment function, or else antivibratory and antinoise functions.